Attitude compensating missile system

ABSTRACT

A missile utilizing directional control guidance concepts in which torques proportional to missile attitude are generated to correct attitude of the missile during its flight to impact at a target.

[451 July 31,1973

llimited States Patent [1 1 McCorkle, Jr.

Edwards et al.

UP 8 57 5666 wwww N4 WH 5 29 5662 .3 2 54 20090 J 2233 S s r I J M n G mN 0 I C H c M E C P m M h m m m W UM TE t I n TS e T m AS I l .l 4 5 5 7Huntsville, Ala.

[73] Assignee: The United States of America as Primary Examiner-VerlinR. Pendegrass Anorney-Harry M. Saragovitz represented by the Secretary0! the Army, Washington, DC.

Apr. 4, 1966 Edward J. Kelly,

Herbert Berl and Harold W. Hilton [22] Filed:

[21] Appl. No.: 540,479

[57] ABSTRACT A missile utilizing directional control guidance con- [52]US. Cl. 244/3 22 [511 Int. F42b 15/02 [58] Field ofSearch................... 244/32, 3.2l, 3.22;

cepts in which torques proportional to missile attitude are generated tocorrect attitude of. the missile during its flight to impact at atarget.

[56] References Cited UNITED STATES PATENTS Goodard. 24413.22 16 Claims,7 Drawing Figures I n I u i PATENTEDJULIH I915 saw Mr 3 William C.McCorkle,Jr.,

INVENTOR. w y f BY EM M #MQW/MV PATENTED I975 3. 749 334 sum 2 OF 3 FIG.3

William C.McCo rk|e,Jr.,

JNVENTUR.

BY 99% MJ/% PATENIED JUL 3 I I975 sum 3 0F 3 William C. McCorkle, Jr.,

INVENTOR. y?

ATTITUDE COMPENSATING MISSILE SYSTEM The device includes a small solidpropellant motor as a rotor gimbaled to provide two degrees of freedom(freedom of motion in pitch and yaw) and a gas deflector system todivert the solid propellant motor exit gas flow to provide missileattitude correction side forces. The motor is tailored to the missileacceleration profile to provide a thrust floated gyro which minimizesgimbal bearing loads and thus reduces the gyro drift rate.

Improvement of free rocket accuracy has long been a major objective inweapons research. Rockets are generally superior to cannons in manyimportant areas, such as rate of fire, payload to launcher (or gun)weight ratio, payload to total logistic weight ratio, range capability,and others. The inability to achieve accurate fire with rockets, withoutthe addition of expensive and generally unreliable guidance systems, hasbeen a major drawback and prevented their more general use as artillery.

The quest for accuracy improvement has, in the last few years, includedthe investigation of simplified guidance schemes. The best known ofthese schemes at present is the so called directional control conceptwhich utilizes an attitude reference provided by a two degree of freedomgyro to obtain missile attitude information used as an input to aconventional control system which generates corrective torquesproportional to missile attitude. The directional control principle isutilized to compensate for the three principle directional errorsources, i.e.; (1) linear thrust malalignment or failure of the thrustaxis to pass through the rocket center of mass by a certain distance;(2) Mallaunch or the unpredictable component of angular motion about atransverse axis at the instant the rocket leaves the launcher; (3) crosswind effect. Stable pockets turn upwind, and the thrust drives therocket off the intended path. The cross wind effect is the mostdifficult error to suppress.

The directional control scheme has been utilized to compensate for theaforementioned sourcesof error; however, the directional control schemeinvolves the use of a good gyro, electrical pickoffs, amplifiers,actuators, power supplies and all the evils and complexity these itemsentail from the standpoint of cost, maintainanceand reliability.

The present invention overcomes the undesirable features noted above byproviding a system wherein the attitude reference and control system arecombined in a single unit thus providing a compact and inexpensivemechanization of the directional control approach which is applicable toa wide range of weapon systems.

The guidance and control system of the present invention includes a twodegree of freedom gyro wheel, spun up by a peripheral solid propellantmotor. The gyro wheel is provided with a chamber loaded with a solidpropellant which is disposed for combustion to provide reaction forcesfor rotation of the gyro wheel.

Upon ignition of the internal solid propellant charge, mass isejectedthrough a nozzle affixed to the wheel and aligned with the spin axisthus providing a space stabilized gas jet. The jet is divided into twoorthogonal planes, turned through an angle of 90 by a fixed deflector,and exits through four ports located on the missile periphery. Thesejets provide the missile attitude control forces. As the missileattitude changes, the deflector moves relative to the space stabilizedgas jet generating corrective torques proportional to the missileattitude. The mass ejected from the gyro wheel produces a thrust forceon the gyro to provide an acceleration thereon.

The present invention provides for accelerations on the gyro, producedby the ejected masses therefrom, which is approximately equal to theaccelerations on the missile, and thus, provide a thrust floated gyrowhich alleviates many of the problems associated with design of gyrogimbals for high G applications.

It is an object of the present invention, therefore, to provide amissile having guidance and control components which are incorporatedinto a single compact unit.

A further object of the present invention is to provide such a guidanceand control unit which is thrust floated" whereby the accelerationsacting on the missile and guidance and control unit are equal.

It is a still further object of the present invention to provide amissile with a solid propellant gyro which is rotatable responsive toignition of the propellant to provide guidance for the missile and toprovide control of the missile by utilizing the thrust produced by thepropellant of the gyro.

A feature of the present invention provides mounting of the gyro in amanner which al.lowsprealignment of the gyros dynamic axis so as togreatly reduce the uncaging transient due to mass unbalance of thewheel.

Further objects, features, and advantages of the present invention willbecome more readily apparent from the following description, taken inconjunction with the accompanying drawings in which:

FIG. I is an elevational view, partially in section of a missileutilizing the principles of my invention.

FIG. 2 is an enlarged elevational view, partially in section, of therocket attitude reference and controller of the present invention.

FIG. 3 is an enlarged elevational view, partially in section, of thegimbals and support structure of the device illustrated in FIG. 2, withthe solid propellant and easing removed for clarity.

FIG. 4 is an enlarged elevational view, partially in section of thegimbal and bearing housing of the gyro of the present invention.

FIG. 5 is a view along lines 5--5 of FIG. 3 illustrating the gimbalarrangement of the gyro assembly.

FIG. 6 is a view along lines 6-6 of FIG. 2 illustrating the gyro spinupnozzle, the gyro propellant, easing, bearing housing and supportstructure being removed for clarity.

FIG. 7 is an elevational view, partially in section, of anotherembodiment of the present invention.

Referring to the drawings and particularly to FIG. I, a missile 10 isprovided with an attitude reference and controlling mechanism I2disposed forward of the center of gravity 13 of the missile. Theattitude reference and controller mechanism 12 includes a small solidpropellant motor 14 as a rotor gimbaled to provide two degrees offreedom (freedom of motion in pitch and yaw) and a gas deflector system16 to divert the solid propellant exit gas flow to provide missileattitude correction side forces. A caging mechanism 18 is connected tothe gyro at the forward end 20 thereof for uncaging or release of thegyro to permit unrestrained rotation thereof.

As shown in FIGS. 2, 3, 4 and 5, mechanism I2 includes a gimbal housing22 having a rotor shaft 24 retatably secured in a bearing assembly 26carried in housing 22. The rotor shaft extends from the bearing assemblyfor rotational support in high speed ball bearings 28 carried in asleeve 30 mounted to a standard two-degree of freedom ball bearinggimbal ring structure 32. Gimbal ring 32 and the bearings are supportedin, a gyro support structure 34 which serves to mount the entiregyroscope assembly in the forward part of the missile structure. Supportstructure 34 includes an annular member 36 having gimbal ring structure32 mounted adjacent the rearward end thereof. The forward end of member34 supports caging mechanism 18. The caging mechanism (FIG. 2) includesa slidable piston 38 having a conical seat 40 thereon for supporttherein of an extension of a caging pin 42 of gimbal housing 22. Piston38 is spring loaded by a compression spring 43 for engagement of thepiston with caging pin 22 to firmly lock the gimbal mechanism in place.A firing squib 44 is mounted in caging mechanism 18 for actuationthereof to permit uncaging of the gyro.

The complete assembly is rigidly mounted in the missile by means of aflange 46 secured to member 36 and extending outwardly therefrom andsecured to an outer annular member 48 carried by the missile frame.

Solid propellant motor 14 is mounted peripherally about member 36 andincludes a propellant 52 carried in chamber 55 of a toroidal casing 54which is secured to the rotor shaft at 56. Casing 54 is provided with aplurality of passages 58 which communicate into the interior of a nozzleassembly 60 secured to the rearward end of casing 54. A charge ofpropellant 64 is carried peripherally about casing 54 and provided withsuitably canted tangential nozzles 66 arranged to form a Heros engine(FIGS. 1 & 6) to set rotor shaft 24 in motion. The propellant is carriedin an annular internally threaded member 63 which is screwed to theforward end of casing 54.

To divert the gases from motor 14 for utilization as control forces onthe missile, a deflector device 16 is rigidly secured to the missileframe and includes a plurality of nozzles or channels 65 communicatingwith ports67 opening through the sides of the missile to the atmosphere.Typically, the deflector device includes four channels or nozzles 65substantially equally spaced and extending radially to communicate withthe atmospherethrough ports 67. The channels are divided by a splitter68 disposed on the centerline of nozzle assembly 60. Withthe gyroaligned with the missile axis, the flow is evenly divided into the 'fourseparate channels. lt is to be understood that the provision of fournozzles is only typical or illustrative and that a plurality of nozzlesmay be used. lt is necessary, that at least three nozzles be used toprovide the restoring torques.

ln operation,spin'-up propellant 64 and gyro propellant 52 are ignitedsimultaneously from an external squib (not shown) and the tire iscarried to the ignitor material by a mild detonating fuze. Immediatelyafter burnout of the spin-out propellant, squib 44 of caging mechanism18 is ignited and piston 38 is pressurized by gas fired by the squib andmoved rapidly from pin 42. A spring loaded detent 43 holds piston 38away from pin 42 responsive to movement of the piston, for uncaging ofthe gyro.

Just prior to uncaging, propellant charge 52 is ignited. The exhaustfrom this combustion is carried through passages 58 to nozzle assembly60 to provide a thrust which would if the rotor were free to move,provide the entire mass carried on gimbal ring structure 32, andincluding approximately one-half of the gimbal ring mass, anacceleration approximately equal to the missile acceleration. Thepurpose of such an acceleration applied to the rotor and gimbal-bornestructure is to minimize the effects of gimbal imperfections andmalalignments, as well as to reduce the gimbal friction by effectivelyunloading the gimbal bearings. The term thrust-floated is thus used todescribe this phenomenon when obtained by the reaction of gasesdischarged from the rotor through a nozzle concentric with the rotorspin axis. In the usual case (non thrust-floated), the center of mass ofthe gimbal borne structure must lie on both gimbal axes (assumedorthogonal and in the same plane). The rotor spin axis is, in this typeof gyroscope, always orthogonal to the plane determined by the gimbalaxes, called the gimbal plane. Location of the center of mass along thespin axis relative to the gimbal plane is not critical due to theassumed direction of missile acceleration relative to the spin axis, (i.e.', essentially along the spin axis) but is critical with respect tocenter of mass location in the gimbal plane. A deviation of the centerof mass in the gimbal plane from the point determined by theintersection of the two gimbal axes (gimbal center), produces, under theassumed missile acceleration condition, a precessing torque proportionalto the acceleration, the gimbalborne mass, and the deviation distance ofthe center of mass from the gimbal center. Such precession is highlyundesirable since the reference axis established by the gyro changes,instead of remaining fixed in space as intended. The thrust-floatedgyroscope presented here, by removing the disturbing torque to theextent that the thrust to gimballed mass ratio matches missileacceleration, either permits greater tolerance on center of mass-gimbalcenter alignment, or greater accuracy (less drift) for a given set ofmanufacturing tolerances. Also, and equally as important, the unloadingof the gimbal bearings achieved by thrust-floating greatly reducesfriction, which, in the presence of gimbal motion (due to missileangular motion), is a significant source of disturbing torques.

The flow from nozzle 60 impinges on a four-way deflector which dividesthe flow into four streams which are then turned out from the missileaxis and exit through the ports in the missile skin placed 90 around themissile. When the nozzle is centered on the deflector axis, the flow isevenly divided among the four streams, and-no net force on the missileresults. However, .when the missile moves transaxially to the rotor spinaxis, the flow is divided unequally to provide a variable flow throughthe nozzles of the deflector assembly, and a net reaction force appearsto drive the missile back into alignment with the gyro axis responsiveto movement of the missile relative to the rotor, as required for thedirectional ontrol guidance concept. The above arrangement assumes theexit ports are forward of the missile center of gravity. If the portsare aft of both the missile center of gravity and the aerodynamic centerof pressure, the flow passages must be arranged so that flow enteringany channel of the deflector exits on the opposite side of the missilefrom that shown in the drawings and discussed above.

The magnitude of the control force F generated in either control planeis given by Where T is the thrust from the rotor, I is distance from thegimbal center to the deflector entrance (or equivalcntly, the nozzleexit plane), R the radius of the nozzle exit, f a factor less than unityexpressing deflector efficiency, and d) the angle between the missile(or deflector axis) and the gyro axis for the control plane in questron.

It is to be understood that other spinning mechanisms may be utilized tospin up the gyro, such as a spring motor, an electric motor, or gas jetsimpinging tangentially on pockets milled into the rotor.

Another embodiment of the present invention is illustrated in FIG. 7wherein like reference numerals refer to like parts. In this embodimentthe spin-up motor is removed from the gyro motor, and spin-up isaccomplished by a separate device which simultaneously provides severalother functions and features, in addition to spin-up including:

a. Uncaging;

b. Covering of spin-up gas escape ports in the missile after completionof spin-up;

0. Missile axial spin after start of missile forward motion;

d. A relatively uncluttered axial passage through the device, desirablefor certain special applications.

As shown in FIG. 7 reference and attitude controller mechanism 12includes a rotor casing having solid propellant 52 therein forming solidpropellant motor 14. Casing 70 is mounted on a rotor shaft 72 which isrotatably supported in a hollow shaft 74 rigidly affixed to the missilethrough an externally threaded sleeve or lead screw 76 which, in turn,is secured to a cylindrical member 78 secured to a support member orframe 80 of the missile. Rotor shaft 72, is mounted for rotation inhollow shaft 74 by means of ball bearing assembly 82 having both innerand outer races 84 and 86, respectively, provided with a sphericalconfiguration.

To provide for spin-up and uncaging of the gyro, a spin up motor 86 ismounted on ball bearings 88 carried on the periphery of lead screw 76and is free to move forward thereon. Spin-up motor 86 includes acylindrical portion 87 enclosing lead screw 76. The spinup motor is heldinitially against rotor casing 70 by a helical spring 90 carried incylindrical support member 78. Spring 90 compresses a plurality of softuncaging flat springs 92 against casing 70. Springs 92 are carried ingrooves 94 provided on the spin-up motor and biased outwardly thereof inengagement with the rotor casing.

Driving connection between the rotor casing and spin-up motor isprovided through a plurality of drive pins 96 mounted in apertures 98and 100 of the motor and rotor casing, respectively.

A lead nut 102 is disposed in threaded relation on lead screw 76 and ismoved thereon by means of a plurality of pins 104 projecting throughslots 106 provided on cylindrical portion 87 of the spin-up motor. Abraking mechanism 108 mounted about cylindrical portion 87 coacts withthe inner ogive surface 110 of the missile and a plurality ofcompression springs 1 12 to arrest motion of the motor and to impartspin to the missile in a manner described hereinbelow.

As shown in FIG. 7, the spin-up motor assembly includes the brakingmechanism 108 having braking shoes 114 provided with a conicalconfiguration. The braking shoes are rigidly secured to a central motorportion 116 provided with a chamber 118 having pro pellant 120 therein.A plurality of spin-up nozzles 122 communicate 'with chamber 118 tocarry combustion gases therefrom and provide rotation of the motor.Combustion gases are exhausted to the atmosphere through a plurality ofports 124 disposed around the periphery of the missile. A conical member126 is secured to central motor portion 116 for slidable movementtherewith to seal spin-up exhaust ports 124.

Operation of the device of this embodiment is initiated upon ignition ofspin-up motor propellant 120 and rotor propellant 118, the gas flowthrough spin-up nozzles 122 drives the rotor up tp speed. As rotation ofthe spin-up motor begins, lead nut 102 advances on lead screw 76 untildrive pins 104 reach the forward end of slots 106, uncaging of the gyrocommences, and by this time, spin-up propellant 120 is completelyconsumed. lnertia of the spin-up motor continues to advance lead nut 102disengaging rotor drive-pins 96, leaving uncaging flat springs 92pressing against the rotor. Further advance of lead nut 102 permitssprings 92 to reach the limit of their extension, and they leave contactwith the rotor, having performed the function of stabilizing the rotorattitude between the time of drive pins 96 disengagement and springs 92disengagement. Further ad Vance of the lead nut brings spin-up motorconical brake shoe 114 into engagement with missile inner ogival surface110, starting compression of springs 112, thereby arresting motion ofthe spin-up motor relative to the missile structure, and transferringthe spin-up motor angular momentum to the missile causing it to beginaxial rotation. The missile propulsion motor is ac tivated just prior toengagement of the brake shoe with the inner missile surface. Atcompletion of spin-up motor braking, the spin exhaust cover member 126seals the spin exhaust ports 124, and the control device activation iscompleted. Gases are expelled through nozzle assembly 16 as described inconjunction with the embodiment described above, for control of missileattitude.

If desired, missile spin could be provided without the use of brake shoe114 and associated structure. For example, threads on the end of leadscrew 76 adjacent cylindrical member 78, may be blunted so that as leadnut i 102 advances on lead screw 76 the nut secured engages the bluntedthreads of lead screw 76 to impart rotational movement to cylindricalmember 78 and the mis sile frame 80.

It should be understood that various specific embodiments disclosed aremerely illustrative of the general invention and that many modificationsthereof may be resorted to that is within the spirit and scope of thepresent invention.

What is claimed is:

1. A directional control system for maintaining a missile on a launchpredetermined path comprising:

a. a gyroscope mounted with its spin axis and rotor aligned with thelongitudinal axis of the missile at launch, said missile moveabletransaxially relative to said rotor;

b. at least three substantially equally spaced and radially alignedports mounted on the missile in communication with the atmosphere;

c. a source of pressurized fluid carried by said gyro rotor forproviding a reaction force on said rotor which is proportional to thereaction force acting on said missile as a result of the thrustdeveloped therein;

d. a jet nozzle disposed in concentric alignment with said spin axes ofsaid gyro rotor and in communication with said ports for directing saidpressurized missile for spin stabilization thereof subsequent to fluidalong the spin axis of said rotor whereby reacrelease of said casing.tion of the expelled pressurized fluid from said 7. A missile as inclaim 6 wherein said gyroscope is rotor through said concentric nozzleprovides a redisposed forwardly of the missile center of gravity. actionforce on said gyro to impart an acceleration 8. A missile as in claim 7wherein said drive means thereto substantially equal to the missileacceleracomprises: tion to provide for unrestricted movement of said a.an annular drive member mounted concentrically gyro. about said hollowshaft for rotation;

e. said gyroscope having said rotor operatively conb. actuating meansfor rotating said annular drive nected between said source and saidports for pro- 10 member; viding a variable fluid flow therethrough toprovide c. braking means carried by said annular drive membalancingforces to the missile responsive to its her for engagementwith the frameof said missile movement relative to said rotor. for imparting said spinto said missile responsive to 2. A missile as in claim 1 wherein saidattitude referrotation of said annular drive member. ence device isdisposed forwardly of the missile center 9. A missile as in claim 8including: of gravity. a. a support member disposed for support of saidhol- 3. A missile as in claim 2 wherein said gyroscope low shaft andextending therefrom in biased relacomprises: tion with the frame of saidmissle; and

a. a bearing housing secured to the frame of said misb. movable meansdisposed for movement along the sile and having a plurality of bearingstherein and a rotor shaft mounted in said bearings and extending fromsaid housing; c. means coacting with said movable means and said b. acasing secured to said rotor shaft for rotation annular drive member forrelease thereof from said therewith and defining said rotor, said casinghav- 2 casing and for moving said brake means into the ing a combustioncharge mounted therein to form engagement with the missile frame forimparting said source of pressurized fluid; spin to said missile.

c. actuating means for initiating rotational movement 10. A missile asin claim 9 wherein said actuating of said casing; means comprises:

d. means for igniting said combustion charge for rotaa. a combustiblecharge carried in an annular chamtion of said casing and said rotorshaft; and, her in said annular drive member and disposed for e. meansfor securing said housing to the frame of ignition to provide propulsivegases for rotation of said missile in gimballed relation thereto. saidmember;

4. A missile as in claim 3 wherein said actuating b. aplurality ofnozzles disposed about the periphery means comprises: 3 5 of saidannular member and in communication a. an annular member disposed aboutthe periphery with said chamber to exhaust said propulsive gases of saidcasing; therefrom for the rotation of said drive member.

b. propellant means carried in a chamber provided in 11. A missile as inclaim 10 wherein said annular member and disposed for ignition to a.said support member is provided with a cylindrical provide propulsivegases for initial rotation of said 40 configuration having externalscrew threads casing; I thereon;

I c. a plurality of nozzles disposed peripherally about b. said meanscoacting with said movable means insaid annular member and incommunication with cludes a nut mounted on said support member in saidchamber to exhaust said propulsive gases threaded relation therewith;therefrom for the rotation of said casing. c. a plurality of pinscarried in said nut and extending 5. A missile as in claim 4 including:into slots provided on the inner. periphery of said a. caging means forsecuring saidhousing against annular drive member for rotation therebyand length of said support member responsive to rotation of said drivemember.

movement prior to initiation of rotational movement of said rotor bysaid actuating means and for movement along said support memberresponsive to rotation of said drive member, said pins disposed releaseof said housing to permit unrestrained for engagement with said annulardrive member removement thereof subsequent to initiation of saidsponsive to a predetermined length of travel of said rotational movementby said actuating means; and, nut on said support member for movement ofsaid b. means for energizing said caging means for release drive memberalong the longitudinal axis of said of said housing. missile for releaseand uncaging of said gyro wheel, 6. A missile as in claim 1 wherein saidgyroscope said pins disposed for continued movement of said comprises:annular drive member until said brake means ena. A hollow shaft securedto the frame of said missile; gages the frame of said missile to impartspin b. a rotor shaft rotatably mounted in said hollow shaft thereto.

in gimballed relation thereto and extending there- 12. A missile as inclaim 11 including:

from; a. a plurality of flat springs mounted on an end surc. a casingsecured to said rotor shaft for rotation face of said annular drivemember and in biased entherewith and defining said rotor, said casinghavgagement with said casing for retention thereof in ing saidcombustion charge therein; the caged position; d. drive means releasablysecured to said casing for b. a plurality of pins mounted in slotsprovided in said surface of said annular drive member for insertion incorrespondingly spaced slots provided on said casing to provide thereleasable connection beimparting initial rotation thereto, said drivemeans disposed for release of said casing for unrestrained movementthereof and for imparting spin to the tween said drive member and saidcasing for the initial rotation thereof.

13. A missile as in claim 12 including:

a. a plurality of ports disposed about the periphery of said missile forexhausting propulsive gases of said drive member actuating means to theatmosphere;

b. a member carried by said annular drive member for movement therewithalong the longitudinal axis of said missile to seal off said ports atcompletion of burning of said combustion charge of said drive member.

14. A missile as in claim 13 wherein said control means comprises:

a. a nozzle secured to said casing in axial alignment with thelongitudinal axis of said missile and said rotor shaft, said nozzlehaving a plurality of ports disposed around the periphery thereof incommunication with the interior of said casing and nozzle; and,

b. said passage means including a plurality of channels secured to theframe of said missile in communication with the exit of said nozzle andthe atmosphere, said channels being concentrically disposed about saidnozzle and joining at a point along the longitudinal axis of saidmissile and said nozzle for equal distribution of said combustion gasesthrough each of said channels when the axes of said reference and saidmissile coincide.

15. A missile disposed for flight in a trajectory ending in a targetcomprising:

a. propulsion means carried by said missile for propelling said missileto a target;

b. mechanism carried by said missile for retention thereof in thetrajectory, said mechanism including a space oriented attitude referencedevice having a gyro wheel provided with a combustion charge thereindisposed for ignition for providing thrust producing gases for rotatingsaid gyro wheel; and, control means including a plurality of passagescommunicating with said gyro wheel and the atmosphere for directing saidthrust producing gases to the atmosphere through predetermined ones ofsaid passages for providing restoring torques on said missile forrestoration of said missile in the trajectory responsive to deviationstherefrom;

c. caging means for securing said gyro wheel against movement prior torotational movement thereof and for release of said gyro wheel to permitunrestrained rotational movemenit thereof, said caging means including ahousing having a piston slidably mounted therein, detent means disposedon said piston for supporting the gyrorotor in a caged position, andmeans for actuating said piston for movement thereof for release of thegyrorotor to an uncaged position to permit unrestrained movementthereof.

16. A missile as in claim 15 wherein said control means comprises:

a. a nozzle secured to said casing in axial alignment with thelongitudinal axis of said missile and said rotor shaft, said nozzlehaving a plurality of ports disposed around the periphery thereof incommunication with the interior of said casing and nozzle; and,

b. said passage means including a plurality of channels secured to theframe of said missile in communication with the exit of said nozzle andthe atmosphere, said channels being coneentrically disposed about saidnozzle and joining at a point along the longitudinal axis of saidmissile and said nozzle for equal distribution of said combustion gasesthrough each of said channels when the axes of said reference device andsaid missile coincide.

1. A directional control system for maintaining a missile on a launchpredetermined path comprising: a. a gyroscope mounted with its spin axisand rotor aligned with the longitudinal axis of the missile at launch,said missile moveable transaxially relative to said rotor; b. at leastthree substantially equally spaced and radially aligned ports mounted onthe missile in communication with the atmosphere; c. a source ofpressurized fluid carried by said gyro rotor for providing a reactionforce on said rotor which is proportional to the reaction force actingon said missile as a result of the thrust developed therein; d. a jetnozzle disposed in concentric alignment with said spin axes of said gyrorotor and in communication with said ports for directing saidpressurized fluid along the spin axis of said rotor whereby reaction ofthe expelled pressurized fluid from said rotor through said concentricnozzle provides a reaction force on said gyro to impart an accelerationthereto substantially equal to the missile acceleration to provide forunrestricted movement of said gyro. e. said gyroscope having said rotoroperatively connected between said source and said ports for providing avariable fluid flow therethrough to provide balancing forces to themissile responsive to its movement relative to said rotor.
 2. A missileas in claim 1 wherein said attitude reference device is disposedforwardly of the missile center of gravity.
 3. A missile as in claim 2wherein said gyroscope comprises: a. a bearing housing secured to theframe of said missile and having a plurality of bearings therein and arotor shaft mounted in said bearings and extending from said housing; b.a casing secured to said rotor shaft for rotation therewith and definingsaid rotor, said casing having a combustion charge mounted therein toform said source of pressurized fluid; c. actuating means for initiatingrotational movement of said casing; d. means for igniting saidcombustion charge for rotation of said casing and said rotor shaft; and,e. means for securing said housing to the frame of said missile ingimballed relation thereto.
 4. A missile as in claim 3 wherein saidactuating means comprises: a. an annular member disposed about theperiphery of said casing; b. propellant means carried in a chamberprovided in said annular member and disposed for ignition to providepropulsive gases for initial rotation of said casing; c. a plurality ofnozzles disposed peripherally about said annular member and incommunication with said chamber to exhaust said propulsive gasestherefrom for the rotation of said casing.
 5. A missile as in claim 4including: a. caging means for securing said housing against movementprior to initiation of rotational movement of said rotor by saidactuating means and for release of said housing to permit unrestrainedmovement thereof subsequent to initiation of said rotational movement bysaid actuating means; and, b. means for energizing said caging means forrelease of said housing.
 6. A missile as in claim 1 wherein saidgyroscope comprises: a. A hOllow shaft secured to the frame of saidmissile; b. a rotor shaft rotatably mounted in said hollow shaft ingimballed relation thereto and extending therefrom; c. a casing securedto said rotor shaft for rotation therewith and defining said rotor, saidcasing having said combustion charge therein; d. drive means releasablysecured to said casing for imparting initial rotation thereto, saiddrive means disposed for release of said casing for unrestrainedmovement thereof and for imparting spin to the missile for spinstabilization thereof subsequent to release of said casing.
 7. A missileas in claim 6 wherein said gyroscope is disposed forwardly of themissile center of gravity.
 8. A missile as in claim 7 wherein said drivemeans comprises: a. an annular drive member mounted concentrically aboutsaid hollow shaft for rotation; b. actuating means for rotating saidannular drive member; c. braking means carried by said annular drivemember for engagement with the frame of said missile for imparting saidspin to said missile responsive to rotation of said annular drivemember.
 9. A missile as in claim 8 including: a. a support memberdisposed for support of said hollow shaft and extending therefrom inbiased relation with the frame of said missle; and b. movable meansdisposed for movement along the length of said support member responsiveto rotation of said drive member. c. means coacting with said movablemeans and said annular drive member for release thereof from said casingand for moving said brake means into the engagement with the missileframe for imparting spin to said missile.
 10. A missile as in claim 9wherein said actuating means comprises: a. a combustible charge carriedin an annular chamber in said annular drive member and disposed forignition to provide propulsive gases for rotation of said member; b. aplurality of nozzles disposed about the periphery of said annular memberand in communication with said chamber to exhaust said propulsive gasestherefrom for the rotation of said drive member.
 11. A missile as inclaim 10 wherein a. said support member is provided with a cylindricalconfiguration having external screw threads thereon; b. said meanscoacting with said movable means includes a nut mounted on said supportmember in threaded relation therewith; c. a plurality of pins carried insaid nut and extending into slots provided on the inner periphery ofsaid annular drive member for rotation thereby and movement along saidsupport member responsive to rotation of said drive member, said pinsdisposed for engagement with said annular drive member responsive to apredetermined length of travel of said nut on said support member formovement of said drive member along the longitudinal axis of saidmissile for release and uncaging of said gyro wheel, said pins disposedfor continued movement of said annular drive member until said brakemeans engages the frame of said missile to impart spin thereto.
 12. Amissile as in claim 11 including: a. a plurality of flat springs mountedon an end surface of said annular drive member and in biased engagementwith said casing for retention thereof in the caged position; b. aplurality of pins mounted in slots provided in said surface of saidannular drive member for insertion in correspondingly spaced slotsprovided on said casing to provide the releasable connection betweensaid drive member and said casing for the initial rotation thereof. 13.A missile as in claim 12 including: a. a plurality of ports disposedabout the periphery of said missile for exhausting propulsive gases ofsaid drive member actuating means to the atmosphere; b. a member carriedby said annular drive member for movement therewith along thelongitudinal axis of said missile to seal off said ports at completionof burning of said combustion charge of said drive member.
 14. A missileas in claim 13 wherein said control means comprises: A. a nozzle securedto said casing in axial alignment with the longitudinal axis of saidmissile and said rotor shaft, said nozzle having a plurality of portsdisposed around the periphery thereof in communication with the interiorof said casing and nozzle; and, b. said passage means including aplurality of channels secured to the frame of said missile incommunication with the exit of said nozzle and the atmosphere, saidchannels being concentrically disposed about said nozzle and joining ata point along the longitudinal axis of said missile and said nozzle forequal distribution of said combustion gases through each of saidchannels when the axes of said reference and said missile coincide. 15.A missile disposed for flight in a trajectory ending in a targetcomprising: a. propulsion means carried by said missile for propellingsaid missile to a target; b. mechanism carried by said missile forretention thereof in the trajectory, said mechanism including a spaceoriented attitude reference device having a gyro wheel provided with acombustion charge therein disposed for ignition for providing thrustproducing gases for rotating said gyro wheel; and, control meansincluding a plurality of passages communicating with said gyro wheel andthe atmosphere for directing said thrust producing gases to theatmosphere through predetermined ones of said passages for providingrestoring torques on said missile for restoration of said missile in thetrajectory responsive to deviations therefrom; c. caging means forsecuring said gyro wheel against movement prior to rotational movementthereof and for release of said gyro wheel to permit unrestrainedrotational movement thereof, said caging means including a housinghaving a piston slidably mounted therein, detent means disposed on saidpiston for supporting the gyrorotor in a caged position, and means foractuating said piston for movement thereof for release of the gyrorotorto an uncaged position to permit unrestrained movement thereof.
 16. Amissile as in claim 15 wherein said control means comprises: a. a nozzlesecured to said casing in axial alignment with the longitudinal axis ofsaid missile and said rotor shaft, said nozzle having a plurality ofports disposed around the periphery thereof in communication with theinterior of said casing and nozzle; and, b. said passage means includinga plurality of channels secured to the frame of said missile incommunication with the exit of said nozzle and the atmosphere, saidchannels being concentrically disposed about said nozzle and joining ata point along the longitudinal axis of said missile and said nozzle forequal distribution of said combustion gases through each of saidchannels when the axes of said reference device and said missilecoincide.